1. Field of the Invention
The present invention relates generally to gas turbine engine, and more specifically a turbine rotor blade with tip rail cooling.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, such as a large frame heavy-duty industrial gas turbine (IGT) engine, a hot gas stream generated in a combustor is passed through a turbine to produce mechanical work. The turbine includes one or more rows or stages of stator vanes and rotor blades that react with the hot gas stream in a progressively decreasing temperature. The efficiency of the turbine—and therefore the engine—can be increased by passing a higher temperature gas stream into the turbine. However, the turbine inlet temperature is limited to the material properties of the turbine, especially the first stage vanes and blades, and an amount of cooling capability for these first stage airfoils.
The first stage rotor blade and stator vanes are exposed to the highest gas stream temperatures, with the temperature gradually decreasing as the gas stream passes through the turbine stages. The first and second stage airfoils (blades and vanes) must be cooled by passing cooling air through internal cooling passages and discharging the cooling air through film cooling holes to provide a blanket layer of cooling air to protect the hot metal surface from the hot gas stream.
High temperature turbine blade tip section heat load is a function of the blade tip leakage flow. A high leakage flow will induce a high heat load onto the blade tip section. Thus, blade tip section sealing and cooling have to be addressed as a single problem. A prior art turbine blade tip design is shown in FIGS. 1-3 and includes a squealer tip rail 11 that extends around the perimeter of the airfoil flush with the airfoil wall to form an inner squealer pocket 12. The main purpose of incorporating the squealer tip in a blade design is to reduce the blade tip leakage and also to provide for improved rubbing capability for the blade. The narrow tip rail 11 provides for a small surface area to rub up against the inner surface of the blade outer air seal (BOAS) that forms the tip gap. Thus, less friction and less heat are developed when the tip rubs.
Traditionally, blade tip cooling is accomplished by drilling holes into the upper extremes of the serpentine coolant passages formed within the body of the blade from both the pressure and suction surfaces near the blade tip edge and the top surface of the squealer cavity. In general, film cooling holes are built in along the airfoil pressure side and suction side tip sections (P/S film holes 15 in FIG. 2 and S/S film holes 16 in FIG. 3) and extend from the leading edge to the trailing edge to provide edge cooling for the blade squealer tip. Also, convective cooling holes 14 also built in along the tip rail 11 at the inner portion of the squealer pocket provide additional cooling for the squealer tip rail. Since the blade tip region is subject to severe secondary flow field, this requires a large number of film cooling holes that requires more cooling flow for cooling the blade tip periphery. FIG. 1 shows the prior art squealer tip cooling arrangement and the secondary hot gas flow migration around the blade tip section. FIG. 2 shows a profile view of the pressure side with tip film cooling holes 15 and FIG. 3 shows the suction side each with tip peripheral film cooling holes 16 for the prior art turbine blade of FIG. 1.
The blade squealer tip rail 11 is subject to heating from three exposed side: 1) heat load from the airfoil hot gas side surface of the tip rail, 2) heat load from the top portion of the tip rail, and 3) heat load from the back side of the tip rail. Cooling of the squealer tip rail by means of discharge row of film cooling holes along the blade pressure side and suction peripheral and conduction through the base region of the squealer pocket becomes insufficient. This is primarily due to the combination of squealer pocket geometry and the interaction of hot gas secondary flow mixing. FIG. 1 shows the secondary flow 17 passing over the blade tip and a vortex flow 18 generated on the blade suction side surface. The effectiveness induced by the pressure film cooling and tip section convective cooling holes become very limited.
FIGS. 4 and 5 show a prior art turbine blade with a tip rail cooling design. A row of pressure side film cooling holes 15 are located on the pressure side wall of the blade and below the pressure side tip rail discharges a film layer of cooling air slightly upward and out onto the surface of the pressure side wall to flow over the pressure side tip rail as seen in FIG. 5. A similar row of suction side film cooling holes 16 is located on the suction side wall. Two tip rail convective cooling holes 14 discharge cooling air into the squealer pocket 12 and produce a vortex flow 19 of the cooling air as represented by the swirling arrows in both FIGS. 4 and 5. The vortex flows 19 follow a path from the upstream most hole 14 in the squealer pocket all the way to the opening in the squealer pocket along the pressure side wall in the trailing edge region of the blade tip as seen in FIG. 5. These two rows of tip rail convective cooling holes 14 are located adjacent to the inner sides of the tip rails. In the FIG. 4 tip rail design of the prior art, the vortex flow develops on the inner sides of both tip rails and travels along the inner side from the leading edge to the trailing edge of the tip pocket. These vortex flows 19 roll along the tip rails 11 from the leading edge toward the trailing edge and mix with the cooling air discharged from the tip floor convection cooling holes 14 and therefore reduce the cooling effectiveness of the backside cooling of the tip rails 11.